Gas turbine engine airfoil with large thickness properties

ABSTRACT

An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation to U.S. application Ser. No.14/728,514 filed Jun. 2, 2015, which claims priority to U.S. ProvisionalApplication No. 62/008,626 which was filed on Jun. 6, 2014 and isincorporated herein by reference.

BACKGROUND

This disclosure relates to a gas turbine engine airfoil. Moreparticularly, this disclosure relates to a gas turbine engine airfoilhaving a large thickness property for use in turbine blades, forexample.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustorsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The thickness of an airfoil designed for turbomachinery applications isan important characteristic. It is often a result of multidisciplinaryconsiderations including aerodynamics, durability, structure and design.However, recent advances in the design of aerodynamicallyhigh-performing, high-pressure turbine blades, particularly at the tip,have caused increased difficulties in the design of blades.

SUMMARY

In one exemplary embodiment, an airfoil for a gas turbine engineincludes an airfoil with pressure and suction sides that are joined atleading and trailing edges. The airfoil extends a span from a support toan end in a radial direction. 0% span and 100% span positionsrespectively correspond to the airfoil at the support and at the end.The leading and trailing edges are spaced apart from one another anaxial chord in an axial direction. A cross-section of the airfoil at aspan location has a diameter tangent to the pressure and suction sides.The diameter corresponds to the largest circle fitting within thecross-section. A ratio of the diameter to the axial chord is at least0.4 between 50% and 95% span location.

In a further embodiment of the above, a root supports a platform thatcorresponds to the support. The end provides a terminal end of theairfoil.

In a further embodiment of any of the above, the root includes a firtree.

In a further embodiment of any of the above, the ratio is at least 0.45,but does not exceed 0.7.

In a further embodiment of any of the above, the ratio is at least 0.5,but does not exceed 0.65.

In a further embodiment of any of the above, the span location isbetween 50% and 70%.

In a further embodiment of any of the above, the ratio of the airfoilpitch to the axial chord is at least 1.6 at 60% span.

In a further embodiment of any of the above, the span location isbetween 75% and 95%.

In a further embodiment of any of the above, the ratio of the airfoilpitch to the axial chord is at least 1.8 at 80% span.

In a further embodiment of any of the above, the airfoil is a turbineblade.

In a further embodiment of any of the above, the airfoil includescooling holes provided on at least one of the leading edge and thepressure side.

In another exemplary embodiment, a gas turbine engine includes acompressor section and a turbine section. A circumferential array ofairfoils is provided in one of the compressor and turbine sections. Thearray has an airfoil that includes pressure and suction sides joined atleading and trailing edges. The airfoil extends a span from a support toan end in a radial direction. 0% span and 100% span positionsrespectively correspond to the exterior wall at the support and the end.The leading and trailing edges are spaced apart from one another anaxial chord in an axial direction. A cross-section of the airfoil at aspan location has a diameter tangent to the pressure and suction sides.The diameter corresponds to the largest circle fitting within thecross-section. A ratio of the diameter to the axial chord is at least0.45 at a 50% span location or greater.

In a further embodiment of the above, the airfoil is provided in theturbine section.

In a further embodiment of any of the above, the airfoil is a turbineblade.

In a further embodiment of any of the above, the ratio is at least 0.45,but does not exceed 0.7.

In a further embodiment of any of the above, the ratio is at least 0.5,but does not exceed 0.65.

In a further embodiment of any of the above, the ratio of the airfoilpitch to the axial chord is between 1.5 and 1.9.

In a further embodiment of any of the above, the ratio of the airfoilpitch to the axial chord is between 1.65 and 1.8.

In a further embodiment of any of the above, the span location isbetween 50% and 70%.

In a further embodiment of any of the above, the span location isbetween 75% and 95%.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2 schematically illustrates a section of the gas turbine engine,such as a turbine section.

FIG. 3 schematically illustrates a turbine blade.

FIG. 4 schematically illustrates a portion of a circumferential array ofblades.

FIG. 5 is a cross-sectional view through an airfoil according to thedisclosure.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis relative to an engine static structure 36 via severalbearing systems 38. It should be understood that various bearing systems38 at various locations may alternatively or additionally be provided,and the location of bearing systems 38 may be varied as appropriate tothe application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axiswhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

Referring to FIG. 2, a cross-sectional view through a high pressureturbine section 54 is illustrated. In the example high pressure turbinesection 54, first and second arrays of circumferentially spaced fixedvanes 60, 62 are axially spaced apart from one another. A first stagearray of circumferentially spaced turbine blades 64, mounted to a rotordisk 68, is arranged axially between the first and second fixed vanearrays. A second stage array of circumferentially spaced turbine blades66 is arranged aft of the second array of fixed vanes 62. It should beunderstood that any number of stages may be used. Moreover, thedisclosed airfoil may be used in a compressor section, turbine sectionand/or fixed or rotating stages.

The turbine blades each include a tip 80 adjacent to a blade outer airseal 70 of a case structure 72, which provides an outer flow path. Thefirst and second stage arrays of turbine vanes and first and secondstage arrays of turbine blades are arranged within a core flow path Cand are operatively connected to a spool 32, for example.

Each blade 64 includes an inner platform 76 respectively defining innerflow path. The platform inner platform 76 supports an airfoil 78 thatextends in a radial direction R, as shown in FIG. 3. It should beunderstood that the turbine vanes may be discrete from one another orarranged in integrated clusters. The airfoil 78 provides leading andtrailing edges 82, 84.

The airfoil 78 is provided between pressure (predominantly concave) andsuction (predominantly convex) sides 94, 96 in an airfoil thicknessdirection (FIG. 5), which is generally perpendicular to a chord-wisedirection provided between the leading and trailing edges 82, 84. Anadditional way of distinguishing between pressure and suction side is tocompare averaged and integrated values of curvature. Between leading andtrailing edges 82, 84, the averaged and integrated curvature is moreconvex on the suction side compared to the pressure side. Multipleturbine blades 64 are arranged in a circumferentially spaced apartmanner in a circumferential direction Y (FIG. 4). The airfoil 78includes multiple film cooling holes 90, 92 respectively schematicallyillustrated on the leading edge 82 and the pressure side 94 (FIG. 4).

The turbine blades 64 are constructed from a high strength, heatresistant material such as a nickel-based or cobalt-based superalloy, orof a high temperature, stress resistant ceramic or composite material.In cooled configurations, internal fluid passages and external coolingapertures provide for a combination of impingement and film cooling.Other cooling approaches may be used such as trip strips, pedestals orother convective cooling techniques. In addition, one or more thermalbarrier coatings, abrasion-resistant coatings or other protectivecoatings may be applied to the turbine vane 64.

FIG. 3 schematically illustrates an airfoil including pressure andsuction sides joined at leading and trailing edges 82, 84. A root 76supports the platform 76. A root 74 may include a fir tree that isreceived in a correspondingly shaped slot in the rotor 68, as is known.The airfoil extends a span from a support, such as an inner platform 76to an end, such as a tip 80 in a radial direction R. The 0% span and the100% span positions respectively correspond to the radial airfoilpositions at the support and the end. The leading and trailing edges 82,84 are spaced apart from one another an axial chord b_(x) (FIG. 5) inthe axial direction X. A cross-section of the airfoil 78, as illustratedin FIG. 5, at a particular span location has a pitch p defined as thecircumferential distance in Y-direction between adjacent airfoils (FIG.4). A cross-section of the airfoil 78, as illustrated in FIG. 5, at aparticular span location also has a diameter d_(max) tangent to thepressure and suction sides 94, 96. The diameter d_(max) corresponds tothe largest circle fitted within the cross-section.

A ratio of the diameter d_(max) to the axial chord b_(x) is at least 0.4between 50% and 95% span location. In one example, the span location isbetween 50% and 70%, and in another example, the span location isbetween 75% and 95%. In another example, the ratio is at least 0.45, butdoes not exceed 0.7. In yet another example, the ratio is at least 0.5,but does not exceed 0.65.

In one example, the ratio of the airfoil pitch to the axial chord isbetween 1.5 and 1.9, and in another example, the ratio of the airfoilpitch to the axial chord is between 1.65 and 1.8.

The cooling holes, such as the cooling holes 90, 92, are provided onvarious locations of the airfoil 78 to provide a boundary layer ofcooling fluid, which protects the airfoil 78 from the hot gases in thecore flow path. The disclosed ratio provides desired heat loadcharacteristics as well as desired film cooling performance. In thepast, such airfoil designs were not considered since they causeincreased weight and pull (for a rotating airfoil). Additionalchallenges are associated with low Mach numbers in the internal coolingpassages. However, airfoils designed with larger thickness near the tip80 may result in overall performance improvement. The larger diameterreduces the pressure gradient near the leading edge, avoiding undesiredflow separation at the leading edge and pressure side.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil for a gas turbine engine comprising:an airfoil including pressure and suction sides joined at leading andtrailing edges, the airfoil extends a span in a radial direction from 0%span position at a support to a 100% span position at an end, theleading and trailing edges are spaced apart from one another an axialchord in an axial direction, a cross-section of the airfoil at a spanlocation has a diameter tangent to the pressure and suction sides, thediameter corresponds to the largest circle fitting within thecross-section, a ratio of the diameter to the axial chord is at least0.45 between 50% and 70% span location wherein the airfoil is configuredto be arranged in a circumferential array of airfoils having an airfoilpitch defined as a circumferential distance between the airfoil and anadjacent airfoil at the 50% span location or greater, and wherein aratio of the airfoil pitch to the axial chord is between 1.5 and 1.9. 2.The airfoil according to claim 1, comprising a root supporting aplatform that corresponds to the support, the end providing a terminalend of the airfoil.
 3. The airfoil according to claim 2, wherein theroot includes a fir tree.
 4. The airfoil according to claim 1, whereinthe airfoil is a rotating airfoil.
 5. The airfoil according to claim 1,wherein the ratio of the diameter to the axial chord does not exceed0.7.
 6. The airfoil according to claim 5, wherein the ratio of thediameter to the axial chord is at least 0.5, but does not exceed 0.65.7. The airfoil according to claim 5, wherein the ratio of the airfoilpitch to the axial chord is at least 1.6 at 60% span.
 8. The airfoilaccording to claim 1, wherein the span location is between 75% and 95%.9. The airfoil according to claim 8, wherein the ratio of the airfoilpitch to the axial chord is at least 1.8 at 80% span.
 10. The airfoilaccording to claim 1, wherein the airfoil is a turbine blade.
 11. Theairfoil according to claim 10, wherein the airfoil includes coolingholes provided on at least one of the leading edge and the pressureside.
 12. The airfoil according to claim 1, wherein the airfoil isunshrouded.
 13. A circumferential array of airfoils, comprising: acircumferential array of airfoils, the array has an airfoil includingpressure and suction sides joined at leading and trailing edges, theairfoil extends a span from a support to an end in a radial direction,0% span and 100% span positions respectively correspond to the exteriorwall at the support and the end, the leading and trailing edges arespaced apart from one another an axial chord in an axial direction, across-section of the airfoil at a span location has a diameter tangentto the pressure and suction sides, the diameter corresponds to thelargest circle fitting within the cross-section, a ratio of the diameterto the axial chord is at least 0.45 at a 50% span location or greater,and an airfoil pitch defined as a circumferential distance betweenadjacent airfoils at the 50% span location or greater, wherein a ratioof the airfoil pitch to the axial chord is between 1.5 and 1.9.
 14. Thecircumferential array of airfoils according to claim 13, wherein thearray of airfoils are rotating airfoils.
 15. The circumferential arrayof airfoils according to claim 13, wherein the ratio of the diameter tothe axial chord does not exceed 0.7.
 16. The circumferential array ofairfoils according to claim 15, wherein the ratio of the diameter to theaxial chord is at least 0.5, but does not exceed 0.65.
 17. Thecircumferential array of airfoils according to claim 13, wherein theratio of the airfoil pitch to the axial chord is between 1.65 and 1.8.18. The circumferential array of airfoils according to claim 13, whereinthe span location is between 50% and 70%.
 19. The circumferential arrayof airfoils according to claim 13, wherein the span location is between75% and 95%.
 20. The circumferential array of airfoils according toclaim 13, wherein the airfoil is unshrouded.